Reversely rotating screw type multiple impeller compressor



A. SABATIUK TATING SCR Sept. 2l, 1954 2,689,681 Ew TYPE MULTIPLE PREssoRl REVERSELY RO 'Filed Sept.

IMPELLER COM 3 Sheets-Sheet 1 @gew 2 Sept 21. 1954 sABATluK A2,689,681

A. REVERSELY ROTATING SCREW TYPE MULTIPLE i Y IMPELLER COMPRESSOR FlledSept. 17, 1949 -3 Sheets-Sheet 2 Eil?, 5.

M/L E7' ABSOLUTE COMPU/VENT RELATIVE E/Y/ Hasan/rf 15x/r Q S550/v0Hora/E eggs-aa? Sept. 21, 1954 A. sABAnUK 2,689,681

REVERSELY ROTATING SCREW TYPE MULTIPLE IMPELLER COMPRESSOR Filed Sept.'17, 1949 3 Sheets-Sheet 3 Eig 5c.

@geni atented S'ept. 2.1., 11954' UNITED STATES REVERSELY ROTATING SCREWTYPE MULTIPLE IMPELLER COMPRESSOR Andrew Sabatiuk, New Britain, Conn.,assignor to United Aircraft Corporation, `East Hartford, Conn., acorporation of Delaware Application September 17, 1949, Serial No.116,286

1 Claim. (Cl. E30- 123) This invention relates to turbo-jet engines andmore specically to axial flow compressors for `such power plants whichprovide high compresling at supersonic speeds relative to the movingrotor blades with the shocks occurring in the passages formed by theblades.

It is another object of this invention to provide an improved compressorfor turbo-jet engines` whereby high compression ratios are produced witha minimum of compression stages and a minimum of space.

Another object of this invention resides in the provision of acompressor which exhausts high pressure fluid of such low relativevelocity that the diffuser can be completely omitted or reduced to aminimum length.

A further object of this invention is to providea multiple stage axialflow compressor wherein each stage includes a rotor operating on theshock-in-rotor principle.

Another object of this invention resides in the employment of the selfstabilizing inlet phenomenon of a supersonic rotor passage therebyeliminating the necessity of statorsbetween rotors in the compressor.

These and other objects will become readily apparent from the followingdescription of the accompanying drawings in which,

Fig. l is a partial cross sectional view of a `turbo-jet engineincluding the compressor form- @and the rotors shown in Fig. 4;

Figs. 5a, 5b and 5c are partial views of the rotor bladesillustratingvarious now conditions through the blades.

Fig. 5d is a detailed view of a rotor blade indicating in detail theiiow through a shock and expansion eld; and

Fig. 6 is a diagrammatic illustration of the compressor pressure ratioand efiiciency `that may be expected of the fluid passing through ashock at various relative inlet Mach numbers.

Although the general practice has been to avoid supersonic relative ilowof compressible uids in compressors due vto-the high loss expec tations,it is known that the supersonic cornlressors are capable of operating atrelatively high compression ratios Without exceeding practicable losslimitations. pressors have utilized the principles wherein supersonicrelative flo-w is obtained in the compressor and the working uid ispassed `through a shock to increase the pressure of the fluid and reduceits velocity below that of sound.

Generally speaking, these supersonic compressors are classied as theshock-in-stator type or the shock-in-rotor type. However, in either easeit has been the practice to utilize stators or turning vanes to obtainproper operation and uid ilow in the rotors. Since the use of stator-sentails a `certain amount of `flow bending, denite losses of varioustypes usually result thereby causing inefciencies which in additionaffect the performance of the rotorblades.

To this end, this invention provides a compressor for a turbo-jet enginewhich compressor in its preferred form operatesy on the shock-in-rotorprinciple but eliminates the usual losses which accompany the use ofstators.

Where compression has been achieved by the shock-in-rotor principle, forexample, the usual practice has been to utilize turning vanes in thecompressor inlet, upstream of the rotor or rotors, to produce. properdirectional l'low and obtain higher relative velocities at the leadingedge ci the rotor blades. With this type of construction the velocity ofthe inlet air ahead of the rotors is adversely limited. In other words,depending upon the shape of the turning vanes and the amount of turningproducedy local supersonic velocity will readily occur over the vanes asincreased subsonic velocities lare approached thereby causingseparation, possible shock and/ which units cooperate to form an airinlet I6V Various types of comand an annular passage I8. The innerhousing I4 carries two concentric shafts 20 and 22 which are rotatablymounted therein by means of a plurality of bearings 24. An axial flowcornpressor is provided and includes in its preferred form a pair ofadjacent rotors 28 and 30 which are xed respectively to the shafts 20and 22. A burner section is provided downstream of the compressorwherein fuel may be injected and the fuel-air mixture ignited to producean expanding high Velocity gaseous medium for rotating the turbineblades 34 and 33 and for further providing a propulsive jet stream inthe exit passage 38. The blades 34 and 36 are operatively connected tothe shafts 22 and 20 respectively whereby the rear turbine rotor drivesthe forward compressor rotor 28 and the forward turbine rotor drives therear compressor rotor 33. The blades 34 and 36 of the adjacent turbinerotors are oppositely inclined relative to the axis of uid flow so thatthe shafts 2D and 22 and their respective compressor rotors 2B and 33are counterrotating. For this reason the compressor rotor blades 5l)and-52 will also be oppositely inclined relative to the axis of flow.The principle operation of this compressor will be more fully describedhereinafter.

Figs. 2 and 3 indicate a modified drive mechanism for interconnectingthe compressor rotors 60 and 62 to the turbine wheel 64. A drive shaft66 is directly connected to the turbine wheel 64 and the forwardcompressor rotor Bl] while the aft compressor rotor 62 is driven in adirection opposite to that of the rotor B0 by means of a planetary gearsystem. This gear system comprises an internally toothed ring gear l0carried by the rotor S2, a pluralityof pinion gears l2 which arerotatably mounted in the xed central housing 16 and a sun gear '18integral with the drive shaft 66. With this construction, then, a singleturbine Wheel may be utilized to rotate both compressor rotors inopposite direction. The various components of the planetary system maybe varied in size as desired to obtain proper relative speeds betweencompressor rotors 6B and 62. f

Where the rotor blades in each rotor are substantially similar in shapeas shown, the rotor 3D, for example, will be rotated at a .somewhatlower speed than the upstream rotor 28 in order to obtain approximatelyidentical supersonic velocities between the blades of both rotors.

Although the iuid leaving the upstream rotor 28 has an absolute velocitywhich is less than the absolute inlet velocity it leaves the rotor 28 ina direction such that due to the rotational Velocity of rotor 3c it hasa relative direction to rotor 3B which is substantially parallel to thechord line of the oncoming blades 52 thereof. Thus the velocity of thefluid entering the rotor til is high relative to the moving rotor blades52 and it is therefore possible to rotate the rotor 30 slower than rotor2S and still maintain substantially the same supersonic relativevelocities through each successive rotor.

It may further be desirable to decrease the pitch of the blades onsuccessive rotors inasmuch as the relative entrance flow to thedownstream rotor is more in line with the chord of the approachingblades.

If for certain reasons a higher Mach number were desirable between therotor blades of the rear rotor 39 than that being obtained through theblading of rotor 28, the aft rotor 30 may be Vrotated at approximatelyVthe same speed. Also,

such identical rotor velocities may be required in the event that theblade structure of each successive rotor varied sufficiently so as tocreate a different flow velocity through the blades. Thus, for example,should a third oppositely rotating rotor stage be used it would bedesirable to have the absolute exit ow from the second rotor to haveadirection more in line with the blades of the third rotor rather thanin the axial direction shown in Fig. 5.

The utilization of successive counter rotating rotors with supersonicflow between the blade passages also provides a self-compensating effectwhich insures that the relative iiow through the blade passages willcontinue to move substantially in line with the chord of the blades andthus insure smoothness and proper directional movement of the exit air.This self-compensating effect is best illustrated by referring to Figs.5a, 5b and 5c which show various conditions of relative ilow into theblades of the second rotor.

Fig. 5a. shows the relative air entering the blades of the second rotorSt substantially chordwise of the blades and represents the optimumoperating condition of this rotor. Fig. 5b illustrates, in exaggeratedform, the direction of relative inlet now when the fluid approaching theblades 52 is such that the blades are at a negative angle of attack. Asthe flow enters the blade passages an expansion field is generated atthe leading edge of the blades adjacent the face thereof. In movingthrough the expansion eld the air flow is deflected toward the face ofthe blades as shown so that it subsequently flows substantially parallelto the face of the blades. At the same time an oblique shock will beproduced near the leading edge on the backside of the blade and thefluid flowing through this shock will also be deflected so that the flowwill continue downstream substantially parallel to the chord of theblades. The bending of flow through ythe expansion eld and the shock ismore clearly shown in detail in Fig. 5d.

Referring to Fig. 5c a condition of iiow opposite to the Fig. 5bsituation is shown wherein the relative inlet iiow to the rotor is suchYthat the approachingrotor blades are relatively at aV high positiveangle of attack.v Under this condition an expansion eld will be producednear the leading edge and on .the backside of each blade while a shockwill be generated adjacent the face of each blade to deflect the flow.Conditions shown in Figs. 5b and 5c will tend to decelerate andaccelerate the inlet axial now respectively so as to approach conditionshown in Fig. 5a.. n Under all the foregoing flow conditions the bladepassages will still generate a normal shock downstream in the vicinityof theA diverging portion of the'pa'ssages. Y

It is to be understood that the same conditions may exist at the bladeentrance of the i'lrst rotor. It is then apparent that with theself-stabilizing effect that the use of stators or turning vanes can becompletely eliminated.

Fig. 6 illustrates the comparsion of theoretical pressure ratio andefficiency at various Mach numbers that may occur when a fluid passesthrough a normal shock which stabilizes in the blades passages of asupersonic compressor. A similar comparative chart may be used as acriterion for choosing design proportions. It will be noted that forexample a pressure ratioof 3 is obtainable in each stage with anefiiciency approximately percent and this pressureA ratio may be furtherincreased at higher Machunumbers with a corresponding decrease inelciency. However, even with some loss of efficiency, at increased Machnumbers within the rotors, extremely high compression ratios areavailable thereby permitting extremely high power output in a turbo-jetengine and the like. It is further apparent that with one stageproducing a compression ratio of 3 the introduction of a second stage ofsimilar theoretical efliciency and without having the normal lossesusually encountered with stators, an overall compression ratio of 9could readily be produced. A compressor of the type described wouldnormally be designed for operation in a region between the lines A and Bwhich are superimposed on the chart of Fig. 6. In other words, the shapeof the blades, the size of the blades, the spacing of the rotors and therotational velocity of the rotors would be of such values that mostefficient operation would be obtained for example in the Mach numberrange between the lines A and B to thereby produce the correspondingcompression ratios and eiliciencies. The compression ratios andefficiencies illustrated in Fig. 6 are by way of example only and arenot limting ranges of this invention.

The axial spacing of the rotors of the compressor of this invention maybe increased somewhat to permit pressure bleed-off and stabilization offlow conditions during starting and offdesign operation. Also the hubtip ratios of the compressor may be altered to obtain optimum passageflow conditions and radial equilibrium.

It is therefore apparent that as a result of this i invention it ispossible to obtain a very high pressure rise across each compressorrotor thereby providing a high pressure operating unit having a highmass flow capacity with minimum space and weight requirements. Such highpressure and high mass flow capacities permit a turbo-jet engine or thelike to operate at unusually high power output with the possible expenseof some losses which are, however, within a practicable limit.

Further, as a result of this invention it is apparent that a highcapacity compressor has been provided which is readily adaptable to agreater number of stages than that illustrated.

It is also apparent that as a result of this invention a compressor andpower plant is provided Which is readily adaptable for aircraft inasmuchas the counterrotating compressor and turbine rotors produce nogyroscopic forces during aircraft maneuvers.

Although certain embodiments of this invention have been illustrated anddescribed herein, it is apparent that various modifications and changesmay be made in the arrangement and construction of the various partswithout departing from the scope of this novel concept.

What it is desired to obtain by Letters Patent is:

A compressor comprising a casing structure having an annularsubstantially unobstructed airflow passage, a plurality of compressorrotors Liournalled axially in said casing structure, blading carried byeach of said rotors and extending across said passage, means forrotating each successive downstream rotor at proportionally lowerrotational speed and in a direction opposite to its preceding rotor,said blades having a thickness which reduces toward their trailing edgesfor at least a distance from mid-chord to said trailing edges therebyforming diverging passages therebetween so as to impart a normal shockto the gaseous medium entering said blades to provide an immediatereduction in the velocity of said medium relative to said rotor bladesand thereby eiiecting an immediate increase in pressure of said medium.

References Cited in the le of this patent UNITED STATES PATENTS

